1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a blade outer air seal for a turbine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
A blade outer air seal (BOAS) is formed from ring segments that are secured by hooks to a ring carrier. The ring segments each include an inner surface that forms a seal with tips of the rotating blades of the turbine. The backside of the ring segments are cooled by directing impingement cooling air onto the surface. To cool the hot side of the ring segment, spent cooling air from the impingement cavity is discharged through film cooling holes to provide for a layer of film cooling air onto the hot surface and to form a seal with the blade tips. Because the inner surface of the ring segment is exposed to the hot gas stream passing through the turbine and the outer surface is exposed to the cool impingement cooling air, a large thermal mismatch is formed within the ring segment.
The above described thermal mismatch that occurs within the ring segments can cause several problems. One is that the ring segments that form the blade outer air seal (BOAS) can warp in the circumferential direction so that a larger leakage flow path is formed in which the hot gas stream can flow. This creates hot spots around the ring segment that can lead to erosion of the material that shortens the useful life of the ring segment.
The thermally induced flexing or warping of the ring segment will also cause any thermal barrier coating (TBC) that is applied to the inner or hot surface to crack or spall off. Missing pieces of the TBC will expose the ring segment metal surface to direct hot gas stream temperature which will also shorten the useful life of the ring segment.